A principal object of the present invention is to provide a low cost system for rendering a space vehicle reusable. There is a need for such a system for use, for example, with space vehicles planned for traveling round trip between a low earth orbit (LEO) and a geosynchronous earth orbit (GEO)--a mission which will involve a change in velocity of approximately 28,670 feet per second during the round trip.
The kinetic energy of a spacecraft during reentry is many times greater than the amount of energy which in terms of heat, would be needed to bring about complete vaporization of the spacecraft. Retardation and friction with the air on reentry into the earth's atmosphere would release a considerable amount of heat and burn up the reentering body. Indeed, this is the normal fate of meteorites entering the atmosphere from outer space.
The "ablating reentry shield" is a known successful way of preventing space vehicle destruction during reentry. Made of ablative material, the shield dissipates heat by melting and vaporizing. Friction with the air heats the ablative material to a temperature of several thousand degrees centigrade, so that the material becomes liquid and "boils off". The zone immediately behind a shock wave which is formed ahead of the returning space vehicle is heated to about 6,000 degrees C. and is in the gaseous state. About 80% of the thermal energy from the intermediate layer between the liquid and the gaseous layer is dissipated as radiation to the surrounding air. The low conductivity of the still-solid ablative material prevents any substantial amount of heat from penetrating into the space vehicle itself during the reentry. (Which is only 200-300 seconds in duration).
Ablative insulation is lightweight and is adequate for the heat load encountered during a reentry, but it requires that the vehicle be returned to the earth after each flight for a very expensive refurbishment.
Many proposals of using a fluid layer for protecting an aircraft and/or reentry vehicle surface against overheating may be found in the patent literature. Examples of such systems are disclosed by the following U.S. Pat. No. 1,426,907, granted Aug. 22, 1922, to George Ramsey; No. 2,468,820, granted May 3, 1949, to Robert H. Goddard; No. 2,995,317, granted Aug. 8, 1961, to Fritz Schoppe; No. 3,026,806, granted Mar. 27, 1962, to Leslie A. Runton and Henry C. Morton; No. 3,062,148, granted Nov. 6, 1962, to John P. Le Bel; No. 3,113,750, granted Dec. 10, 1963, to Melville W. Beardsley; No. 3,259,065, granted July 5, 1966, to Don H. Ross and Eugene S. Rubin; No. 3,298,637, granted Jan. 17, 1967, to Shao-Tang Lee; No. 3,508,724, granted Apr. 28, 1970, to Stanley H. Scher and James C. Dunavant; No. 3,624,751, granted Nov. 30, 1971, to Ronald F. Dettling; No. 3,731,893, granted May 8, 1973, to Charles J. Stalmach, Jr.; No. 3,785,591, granted Jan. 15, 1974, to Charles J. Stalmach Jr.; and No. 4,014,485, granted Mar. 29, 1977, to Laird D. Kinnaird and Seth B. Moorhead, Jr.
The round trip between a GEO and an LEO is an extremely difficult mission. However, the recent successful beginning of the space shuttle program has made it clear that in the near future there will be a need to accomplish this mission frequently and at a low cost. The already difficult mission will be further complicated when manned space vehicles make the round trip.
Various braking techniques have been proposed to achieve the desired change in velocity when a space vehicle returns from a GEO to an LEO. These include using the vehicle's propulsive power for braking, and grazing the upper atmosphere so that aerodynamic drag, or friction with the atmosphere, will convert the vehicle's kinetic energy into heat. Such a grazing maneuver must be done in a precise manner to avoid the loss of too much velocity, which would cause the vehicle to enter the lower atmosphere and burn up, and the loss of too little velocity, which would cause the vehicle to coast up into the Van Allen radiation belt. The difficulty of making the maneuver precise is greatly increased by navigational inaccuracies and unpredictable variations in the atmosphere. The vehicle could be required to correct for variations in atmospheric density of up to 50%. One method of providing such correction is aerodynamic maneuvering, which uses vehicle lift to climb or descent to compensate for density variations.
Both propulsive braking and aerodynamic maneuvering require major changes in the basic structure of the space vehicle. Propulsive braking requires a large increase in propulsive capability, which necessarily involves the addition of propellant and tankage to the detriment of payload. Aerodynamic maneuvering requires aerodynamic control surfaces, center of gravity offset, or similar schemes and complex control and thermal protection systems. The additional equipment in each case significantly increases the cost, complexity, and weight of the vehicle. The increased weight decreases the payload weight to vehicle weight ratio and thereby further increases the cost of the mission.
The problems presented by such increases in vehicle weight and complexity are further aggravated by the weight and size limitations of the space shuttle and the practical limitations of the space vehicle's structure and propulsive capabilities. Many solutions have been proposed, most of which are very complicated and expensive in that they employ multiple shuttle launches and "on orbit" assembly of a "wagon train" of fuel tanks, propulsion systems, and payload. These complicated procedures probably would work, but to be cost effective, each manned mission would have to accomplish a number of separate tasks and, therefore, would have to spend long periods of time both in GEO, to accomplish the tasks, and in LEO, to organize the mission. The necessity for such time-consuming activity would severely reduce the flexibility of these procedures and might even render them totally impracticable.
A possible alternative to the above complicated procedures is a dramatic increase in the initial launch to low earth orbit capability. However, this alternative would be quite expensive and would require extensive research and development.
Another possible alternative is to temporarily increase the vehicle's drag when it approaches an LEO from a GEO so that the desired reduction in velocity can be achieved by grazing the upper atmosphere. The use of balloon-like structures to provide drag and decelerate an object's descent to earth is well known. Examples of such structures are disclosed in U.S. Pat. No. 3,053,476, granted Sept. 11, 1962, to J. L. Mohar, U.S. Pat. No. 3,286,951, granted Nov. 22, 1966, to R. T. Kendall; and U.S. Pat. No. 3,508,724, granted Apr. 28, 1970, to S. H. Scher et al. Many of these and other known structures and procedures work well in slowing a descent to the earth's surface, but they lack the heat protection and control capabilities necessary for an aerobraking maneuver in the upper atmosphere. Moreover, the addition of known methods of cooling would increase the weight and complexity of such structures to a degree that would make them as impractical as propulsive braking and aerodynamic maneuvering.
The above patents and the prior art that is discussed and/or cited therein should be studied for the purpose of putting the present invention into proper perspective relative to the prior art.